Variable shape inlet section for a nacelle assembly of a gas turbine engine

ABSTRACT

A method of improving the aerodynamic performance of an inlet section of a nacelle assembly of a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, (a) sensing an operability condition, (b) adjusting a leading edge of the inlet section in response to the step (a), and (c) adjusting a thickness of the inlet section in response to the step (a).

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. patent application Ser. No.11/769,749, which was filed on Jun. 28, 2007.

BACKGROUND

This disclosure generally relates to a gas turbine engine, and moreparticularly to a gas turbine engine having a variable shape inletsection.

In an aircraft gas turbine engine, such as a turbofan engine, air ispressurized in a compressor and mixed with fuel in a combustor forgenerating hot combustion gases. The hot combustion gases flowdownstream through turbine stages which extract energy from the hotcombustion gases. A fan section supplies air to the compressor.

Combustion gases are discharged from the turbofan engine through a coreexhaust nozzle and a quantity of fan air is discharged through anannular fan exhaust nozzle defined at least partially by a nacelleassembly surrounding the core engine. A majority of propulsion thrust isprovided by the pressurized fan air which is discharged through the fanexhaust nozzle, while the remaining thrust is provided by the combustiongases discharged through the core exhaust nozzle.

It is known in the field of aircraft gas turbine engines that theperformance of a turbofan engine varies during diversified operabilityconditions experienced by the aircraft. An inlet lip section located atthe foremost end of the turbofan nacelle assembly is typically designedto enable operation of the turbofan engine and reduce separation ofairflow from the internal and external flow surfaces of the inlet lipsection during these diversified conditions. For example, the nacelleassembly requires a “thick” inlet lip section to support operation ofthe engine during specific flight conditions, such as crosswindconditions, take-off conditions and the like. Disadvantageously, the“thick” inlet lip section may reduce the efficiency of the turbofanengine during normal cruise conditions of the aircraft, for example. Asa result, the maximum diameter of the nacelle assembly is approximately10-20% larger than required during cruise conditions. Since aircrafttypically operate in cruise conditions for extended periods, turbofanefficiency gains can lead to substantially reduced fuel burn andemissions.

Accordingly, it is desirable to provide a nacelle assembly having anadaptive structure to improve the performance of a turbofan gas turbineengine during diversified operability conditions.

SUMMARY

A method of improving the aerodynamic performance of an inlet section ofa nacelle assembly of a gas turbine engine, according to an exemplaryaspect of the present disclosure includes, among other things, sensingan operability condition, adjusting a leading edge of the inlet sectionin response to the step of sensing, and adjusting a thickness of theinlet section in response to the step of sensing.

In a further non-limiting embodiment of the foregoing method, theleading edge is adjustable between a thin position and a blunt positionand the step of adjusting the leading edge includes the step of movingthe leading edge between the thin position and the blunt position.

In a further non-limiting embodiment of either of the foregoing methods,the thickness of the inlet section is adjustable between a firstposition and a second position and the step of adjusting the thicknessincludes adjusting the thickness between the first position and thesecond position.

In a further non-limiting embodiment of any of the foregoing methods,the second position is radially outward from the first position.

In a further non-limiting embodiment of any of the foregoing methods,the second position is radially inward from the first position.

In a further non-limiting embodiment of any of the foregoing methods,the inlet section includes a plurality of discrete sectionscircumferentially disposed about an engine longitudinal centerline axisand each having a leading edge and a body panel portion aft of theleading edge that includes an adaptive structure. The step of adjustingthe thickness includes adjusting a thickness of each of the body panelportions in each of a radially outer direction and a radially innerdirection relative to the engine longitudinal centerline axis to alterthe adaptive structure.

In a further non-limiting embodiment of any of the foregoing methods,the adaptive structure of a first discrete section of the plurality ofdiscrete sections is altered independently of the adaptive structure ofa second discrete section of the plurality of discrete sections.

In a further non-limiting embodiment of any of the foregoing methods,the step of sensing is performed using a programmable controller.

A method of influencing an adaptive structure of a plurality of discretesections of a gas turbine engine according to another exemplary aspectof the present disclosure includes, among other things, positioning theplurality of discrete sections circumferentially about an enginelongitudinal centerline axis. Each of the plurality of discrete sectionshave a leading edge and a body panel portion aft of the leading edgethat includes the adaptive structure. The method includes adjusting athickness of at least one body panel portion in each of a radially outerdirection and a radially inner direction relative to the enginelongitudinal centerline axis to alter the adaptive structure.

In a further non-limiting embodiment of the foregoing method, the methodincludes the step of identifying an operability condition prior to thestep of adjusting.

In a further non-limiting embodiment of either of the foregoing methods,the method includes the step of moving the leading edge between a thinposition and a blunt position.

In a further non-limiting embodiment of any of the foregoing methods,the adaptive structure of a first discrete section of the plurality ofdiscrete sections is altered independently of the adaptive structure ofa second discrete section of the plurality of discrete sections.

In a further non-limiting embodiment of any of the foregoing methods,the thickness of the first discrete section and the thickness of thesecond discrete section are adjusted uniformly.

In a further non-limiting embodiment of any of the foregoing methods,the thickness of the first discrete section and the thickness of thesecond discrete section are adjusted by a different thickness amount.

In a further non-limiting embodiment of any of the foregoing methods,the plurality of discrete sections include at least a first discretesection and a second discrete section, and the step of adjusting thethickness is performed on one of the first discrete section and thesecond discrete section but not on the other of the first discretesection and the second discrete section.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a general sectional view of a gas turbine engine;

FIG. 2 illustrates a nacelle assembly of a gas turbine engineillustrated in FIG. 1;

FIG. 3 illustrates a general perspective view of the nacelle assembly ofa gas turbine engine shown in FIG. 1;

FIG. 4A illustrates a first example position of a leading edge of aninlet section of the nacelle assembly;

FIG. 4B illustrates a second example position of the leading edge of theinlet section of the nacelle assembly;

FIG. 5 illustrates an example mechanism for manipulating an adaptivestructure of an inlet section of a nacelle assembly; and

FIG. 6 illustrates a side view of the inlet section of the nacelleassembly of a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 which includes (in serialflow communication) a fan section 14, a low pressure compressor 15, ahigh pressure compressor 16, a combustor 18, a high pressure turbine 20and a low pressure turbine 22. During operation, air is pulled into thegas turbine engine 10 by the fan section 14, pressurized by thecompressors 15, 16 and is mixed with fuel and burned in a combustor 18.Hot combustion gases generated within the combustor 18 flow through thehigh and low pressure turbines 20, 22 which extract energy from the hotcombustion gases.

In a two-spool gas turbine engine architecture, the high pressureturbine 20 utilizes the energy extracted from the hot combustion gasesto power the high pressure compressor 16 through a high speed shaft 19,and the low pressure turbine 22 utilizes the energy extracted from thehot combustion gases to power the low pressure compressor 15 and the fansection 14 though a low speed shaft 21. However, the invention is notlimited to the two-spool gas turbine engine architecture described andmay be used with other architectures, such as a single-spool axialdesign, a three-spool axial design and other architectures. That is, thepresent invention is applicable to any gas turbine engine, and to anyapplication.

The example gas turbine engine 10 is in the form of a high bypass ratioturbofan engine mounted within a nacelle assembly 26, in which asignificant amount of air pressurized by the fan section 14 bypasses thecore engine 39 for the generation of propulsion thrust. The nacelleassembly 26 partially surrounds an engine casing 31 that houses the coreengine 39 and its components. The airflow entering the fan section 14may bypass the core engine 39 via a fan bypass passage 30 which extendsbetween the nacelle assembly 26 and the engine casing 31 for receivingand communicating a discharge airflow F1. The high bypass flowarrangement provides a significant amount of thrust for powering theaircraft.

The engine 10 may include a geartrain 23 that controls the speed of therotating fan section 14. The geartrain 23 can be any known gear system,such as a planetary gear system with orbiting planet gears, a planetarygear system with non-orbiting planet gears or other type of gear system.In the disclosed example, the geartrain 23 has a constant gear ratio. Itshould be understood, however, that the above parameters are onlyexamples of a contemplated geared turbofan engine 10. That is, theinvention is applicable to traditional turbofan engines as well as otherengine architectures.

The discharge airflow F1 is discharged from the engine 10 through a fanexhaust nozzle 33. Core exhaust gases C are discharged from the coreengine 39 through a core exhaust nozzle 32 disposed between the enginecasing 31 and a center plug 34 disposed coaxially around a longitudinalcenterline axis A of the gas turbine engine 10.

FIG. 2 illustrates an example inlet lip section 38 of the nacelleassembly 26. The inlet lip section 38 is positioned near a forwardsegment 29 of the nacelle assembly 26. A boundary layer 35 is associatedwith inlet lip section 38. The boundary layer 35 represents an areaadjacent to each of an inner and outer flow surface of the inlet lipsection 38 at which the velocity gradient of airflow is zero. That is,the velocity profile of oncoming airflow F2 goes from a free stream awayfrom the boundary layer 35 to near zero at the boundary layer 35 due tofriction forces that occur as the oncoming airflow F2 passes over theouter and inner flow surfaces of the inlet lip section 38.

The inlet lip section 38 defines a contraction ratio. The contractionratio represents a relative thickness of the inlet lip section 38 of thenacelle assembly 26 and is represented by the ratio of a highlight areaH_(a) (ring shaped area defined by highlight diameter D_(h)) and athroat area T_(a) (ring shaped area defined by throat diameter D_(r)).Currently industry standards typically require a contraction ratio ofapproximately 1.33 to reduce the separation of oncoming airflow F2 fromthe outer and inner flow surfaces of the inlet lip section 38 duringengine operation, but other contraction ratios may be feasible. “Thick”inlet lip section designs, which are associated with large contractionratios, increase the maximum diameter D_(max) and increase the weightand drag penalties associated with the nacelle assembly 26. In addition,a desired ratio of the maximum diameter Dmax relative to the highlightdiameter D_(h) is typically less than or equal to about 1.5, forexample. A person of ordinary skill in the art would understand thatother ratios of the maximum diameter Dmax relative to the highlightdiameter D_(h) are possible and will vary depending upon design specificparameters.

Referring to FIG. 3, the inlet lip section 38 includes a plurality ofdiscrete sections 40 disposed circumferentially about the enginelongitudinal centerline axis A. Each of the discrete sections 40includes a leading edge 42 and a body panel portion 44. Each discretesection 40 has an adaptive structure that is capable of a shape change.The inlet lip section 38 is sectioned into the plurality of discretesections 40 to reduce the stiffness of the closed annular structure ofthe inlet lip section 38 and allow flexure thereof. Each discretesection 40 is designed to be capable of deformation (i.e., the materialsremain within their elastic limits), yet simultaneously have therequisite stiffness to maintain a deformed shape while under aerodynamicand external pressure loads. In addition, as would be understood bythose of ordinary skill in the art having the benefit of thisdisclosure, each discrete section 40 could slightly overlap withadjacent discrete sections 40 to allow the shape change of the inlet lipsection 38 to occur without interference. A fixed nacelle portion 41 ispositioned downstream from the inlet lip section 38.

In one example, the discrete sections 40 are comprised of an aluminumalloy. In another example, the discrete sections are comprised of atitanium alloy. It should be understood that any deformable material maybe utilized to form the discrete sections 40. A person of ordinary skillin the art having the benefit of this description would be able tochoose an appropriate material for the example discrete sections 40 ofthe inlet lip section 38.

Influencing the adaptive structure of the inlet lip section 38 duringspecific flight conditions to achieve a desired shape change increasesthe amount of airflow communicated through the gas turbine engine 10 andreduces the internal and external drag experienced by the inlet lipsection 38. In one example, the adaptive structure of the inlet lipsection 38 is influenced by adjusting the shape of the leading edge 42of each discrete section 40 (see FIGS. 4A and 4B). In another example,the adaptive structure of the inlet lip section 38 is influenced byadjusting a thickness of the body panel portions 44 of each discretesection 40 (see FIG. 5). In yet another example, the adaptive structureof the inlet lip section 38 is influenced by adjusting both the leadingedge 42 and the thickness of the body panel portion 44 of each discretesection 40.

FIGS. 4A and 4B illustrate the adjustment of the leading edge 42 of adiscrete section 40 of the inlet lip section 38 between a first positionX (see FIG. 4A) and a second position X′ (see FIG. 4B). The firstposition X represents a “thin” inlet lip section 38. The second positionX′ represents a “blunt” inlet lip section 38. Each leading edge 42 ismoved between the first position X and the second position X′ via arotary actuator 46, for example. The rotary actuator 46 rotates ineither a clockwise or counterclockwise direction to move a linkageassembly 48 and adjust the leading edge 42 between the first position Xand the second position X′. The rotary actuator 46 and the linkageassembly 48 are mounted within a cavity 50 of each discrete section 40.

At least one linkage assembly 48 is provided within each discretesection 40 and includes a plurality of linkage arms 52 and a pluralityof pivot points 54. The rotary actuator 46 pivots, toggles, extendsand/or flexes the linkage arms 52 of the linkage assembly 48 about thepivot points 54 to move the leading edge 42 between the “thin”, firstposition X and the “blunt”, second position X′. Although the presentexample is illustrated with a rotary actuator and linkage arms connectedvia pivot points, other mechanisms may be utilized to move the leadingedges 42 of the discrete sections 40 between the first position X andthe second position X′, including but not limited to linear actuators,bell cranks, etc. A person of ordinary skill in the art having thebenefit of this disclosure will be able to implement an appropriateactuator assembly to manipulate the leading edge 42 of each discretesection 40. In addition, it should be understood that the leading edge42 is moveable to any position between the first position X and secondposition X′.

The adaptive structure of the inlet lip section 38 is influenced bymoving the leading edge 42 of each discrete section 40 between the firstposition X and the second position X′ in response to detecting anoperability condition of the gas turbine engine 10. In one example, theoperability condition includes a take-off condition. In another example,the operability condition includes a climb condition. In yet anotherexample, the operability condition includes a landing condition. Instill another example, the operability condition includes a high angleof attack condition. It should be understood that the adaptive structureof the inlet lip section 38 is adjustable in response to any operabilitycondition experienced by the aircraft. Each leading edge 42 ispositioned at/returned to the first position X during normal cruiseconditions of the aircraft.

A sensor 61 detects the operability condition and communicates with acontroller 62 to translate the leading edge 42 between the firstposition X and the second position X′ and influence the adaptivestructure of the inlet lip section 38. Of course, this view is highlyschematic. In addition, the illustrations of the movement of the inletlip section 38 are shown exaggerated to better illustrate the adaptivestructure thereof. A person of ordinary skill in the art wouldunderstand the distances the leading edge 42 should be moved between theposition X and the second position X′ in response to sensing a specificoperability condition.

It should be understood that the sensor 61 and the controller 62 may beprogrammed to detect any known operability condition and that eachoperability condition may be associated with a distinct position of theleading edge 42 of the inlet lip section 38. That is, the sensor 61 andthe controller 62 are operable to situate the leading edge 42 of eachdiscrete section 40 at a position which corresponds to the operabilitycondition that is detected. Also, the sensor can be replaced by anycontroller associated with the gas turbine engine 10 or an associatedaircraft. In fact, the controller 62 itself can include the “sensor” andgenerate the signal to adjust the contour of the inlet lip section 38.

FIG. 5 illustrates the adjustment of a thickness T of a body panelportion 44 of each discrete section 40 to influence the adaptivestructure of the inlet lip section 38. The thickness T of the body panel44 is adjustable between a “thin” inlet lip section 38 and a “thick”inlet lip section 38, for example. An inner surface 70 and an outersurface 72 of each body panel portion 44 are moveable in a Y direction(i.e., radially outward) to adjust each discrete section 40 to a “thick”position. In addition, the inner and outer surfaces 70, 72 are moveablein a Z direction to adjust each discrete section to a “thin” position.

The thickness T adjustment of each body panel portion 44 is achieved viaa linear actuator 56 and a linkage assembly 58. The linear actuator 56and the linkage assembly 58 are received in the cavity 50 of eachdiscrete section 40. Although the present example is illustrated with alinear actuator and linkage arms connected via pivot points, othermechanisms may be utilized to adjust the thickness T of each body panelportion 44.

The linear actuator 56 includes an actuator arm 60 which is moveable ina R or L direction to move the linkage assembly 58 and thereby adjustthe thickness of the body panel portion 44. The linkage assembly 58includes a plurality of linkages 64 and a plurality of pivot points 66.The linear actuator 56 adjusts the thickness T of each body panelportion 44 by retracting, pivoting, toggling, extending and/or flexingthe linkages 64 about each pivot point 66. In one example, the actuatorarm 60 of the linear actuator 56 moves in a R direction to retract theouter skin (i.e., move the outer skin in the Z direction) of the bodypanel portion 44 and provide a “thin” inlet lip section 38. In anotherexample, the actuator arm 60 of the linear actuator 56 is moved in a Ldirection to expand the outer skin (i.e., move the outer skin in the Ydirection) of the body panel portion 44 and provide a “thick” inlet lipsection 38. That is, the thickness T of each body panel portion 44 isadjusted either radially outwardly or radially inwardly to provide a“thick” inlet lip section or a “thin” inlet lip section, respectively.

The thickness of each discrete section 40 is adjusted in response todetecting an operability condition. In one example, the operabilitycondition includes a take-off condition. In another example, theoperating condition includes a climb condition. In another example, theoperability condition includes a high angle of attack condition. Instill another example, the operability condition includes a landingcondition. It should be understood that the thickness of the body panelportion 44 may be adjusted to influence the adaptive structure of theinlet lip section 38 in response to any operability conditionexperienced by the aircraft. The thickness T is adjusted/returned to a“thin” position at cruise conditions of the aircraft.

A sensor 61, as is shown in FIGS. 4 a and 4 b, detects the operabilitycondition and communicates with a controller 62 to adjust the thicknessT of each discrete section 40. Of course, this view is highly schematic.In addition, the illustrations of the movement of the inlet lip section38 are shown exaggerated to better illustrate the adaptive structurethereof. A person of ordinary skill in the art would understand thedistances the thickness T should be adjusted in response to sensing aspecific operability condition.

It should be understood that the sensor 61 and the controller 62 may beprogrammed to detect any known operability condition and that eachoperability condition may be associated with a distinct thickness T ofthe body panel portions 44 of the discrete sections 40. That is, thesensor 61 and the controller 62 are operable to adjust the thickness Tof each discrete section 40 to a position which corresponds to theoperability condition that is detected. The thickness T of each discretesection 40 may be adjusted uniformly or differently about thecircumference. In some instances, such as operating during strongcross-winds, for example, only certain discrete sections 40 may beadjusted, while other discrete sections 40 are left unchanged. Also, thesensor can be replaced by any controller associated with the gas turbineengine 10 or an associated aircraft. In fact, the controller 62 itselfcan generate the signal to adjust the contour of the inlet lip section38.

Although illustrated in FIGS. 4 and 5 as having only a single mechanismfor adjusting the shape of the inlet lip section 38 (i.e., one of arotary actuator 46 with a linkage assembly 48 or a linear actuator 56with a linkage assembly 58), it should be understood that each discretesection 40 could include both types of mechanisms to achieve both aleading edge adjustment and a thickness adjustment of the inlet lipsection 38. A person of ordinary skill in the art having the benefit ofthis disclosure would be able to design the inlet lip section 38 toachieve a desired aerodynamic performance level.

Influencing the adaptive structure of the inlet lip section 38 may alsobe achieved during diverse operating conditions by “drooping” a portionof the inlet lip section 38 relative to a remaining portion of the inletlip section 38 (See FIG. 6). In one example, a portion of the discretesections 40 positioned near a top portion 80 of the inlet lip section 38are translated in an X direction and a portion of discrete sections 40positioned near a bottom portion 82 of the inlet lip section 38 aretranslated in a Y direction to create a droop angle D relative to aplane 84 defined by the foremost end 86 of the inlet lip section 38. Thetranslations of the discrete sections 40 in the X and Y directions areachieved via adjustment of linkage assembly 48 (See FIGS. 4 a and 4 b),the linkage assembly 58 (See FIG. 5) or a combination of both thelinkage assembly 48 and the linkage assembly 58. The droop angle D isbetween 2 to 6 degrees relative to the plane 84, in one example.Although FIG. 6 illustrates the “droop” of the bottom portion 82relative to the remaining portion of the inlet lip section 38, it shouldbe understood that any portion of the inlet lip section 38 may bedrooped to improve the aircraft engine performance and reduce nacelledrag at all flight conditions.

The adaptive inlet lip section 38 improves aerodynamic performance ofthe gas turbine engine 10 during all operability conditions experiencedby the aircraft. In addition, because of the shape changing capabilitiesof the inlet lip section 38, the aircraft may be designed having a“thin” inlet lip section 38 (i.e., a slim line nacelle having a reducedcontraction ratio is achieved). As a result, the nacelle assembly 26 isdesigned for specific cruise conditions of the aircraft. A reducedmaximum diameter of the nacelle assembly 26 may therefore be achievedwhile reducing weight, reducing drag, reducing fuel burn and increasingthe overall efficiency of the gas turbine engine 10.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that certain modifications would come within the scope of thisdisclosure. For that reason, the following claims should be studied todetermine the true scope and content of this invention.

What is claimed is:
 1. A method of improving the aerodynamic performanceof an inlet section of a nacelle assembly of a gas turbine engine,comprising the steps of: (a) sensing an operability condition; (b)adjusting a leading edge of the inlet section in response to said step(a); and (c) adjusting a thickness of the inlet section in response tosaid step (a).
 2. The method as recited in claim 1, wherein the leadingedge is adjustable between a thin position and a blunt position and saidstep (b) includes the step of: moving the leading edge between the thinposition and the blunt position.
 3. The method as recited in claim 1,wherein the thickness of the inlet section is adjustable between a firstposition and a second position and said step (c) includes the step of:adjusting the thickness between the first position and the secondposition.
 4. The method as recited in claim 3, wherein the secondposition is radially outward from the first position.
 5. The method asrecited in claim 3, wherein the second position is radially inward fromthe first position.
 6. The method as recited in claim 1, wherein theinlet section includes a plurality of discrete sectionscircumferentially disposed about an engine longitudinal centerline axisand each having a leading edge and a body panel portion aft of theleading edge that includes an adaptive structure, wherein said step (c)includes adjusting a thickness of each of the body panel portions ineach of a radially outer direction and a radially inner directionrelative to the engine longitudinal centerline axis to alter theadaptive structure.
 7. The method as recited in claim 6, wherein theadaptive structure of a first discrete section of the plurality ofdiscrete sections is altered independently of the adaptive structure ofa second discrete section of the plurality of discrete sections.
 8. Themethod as recited in claim 1, wherein said step (a) is performed using aprogrammable controller.
 9. A method of influencing an adaptivestructure of a plurality of discrete sections of a gas turbine engine,comprising the steps of: positioning the plurality of discrete sectionscircumferentially about an engine longitudinal centerline axis, each ofthe plurality of discrete sections having a leading edge and a bodypanel portion aft of the leading edge that includes the adaptivestructure; and adjusting a thickness of at least one body panel portionin each of a radially outer direction and a radially inner directionrelative to the engine longitudinal centerline axis to alter theadaptive structure.
 10. The method as recited in claim 9, comprising thestep of identifying an operability condition prior to the step ofadjusting.
 11. The method as recited in claim 9, comprising the step ofmoving the leading edge between a thin position and a blunt position.12. The method as recited in claim 9, wherein the adaptive structure ofa first discrete section of the plurality of discrete sections isaltered independently of the adaptive structure of a second discretesection of the plurality of discrete sections.
 13. The method as recitedin claim 12, wherein the thickness of the first discrete section and thethickness of the second discrete section are adjusted uniformly.
 14. Themethod as recited in claim 12, wherein the thickness of the firstdiscrete section and the thickness of the second discrete section areadjusted by a different thickness amount.
 15. The method as recited inclaim 9, wherein the plurality of discrete sections include at least afirst discrete section and a second discrete section, and said step ofadjusting the thickness is performed on one of the first discretesection and the second discrete section but not on the other of thefirst discrete section and the second discrete section.